The present application relates generally to combustion turbine engines, which, as used herein and unless specifically stated otherwise, includes all types of combustion turbine engines, such as those used in power generation and aircraft engines. More specifically, but not by way of limitation, the present application relates to apparatus, systems and/or methods for cooling the platform region of turbine rotor blades.
A gas turbine engine typically includes a compressor, a combustor, and a turbine. The compressor and turbine generally include rows of airfoils or blades that are axially stacked in stages. Each stage typically includes a row of circumferentially spaced stator blades, which are fixed, and a set of circumferentially spaced rotor blades, which rotate about a central axis or shaft. In operation, the rotor blades in the compressor are rotated about the shaft to compress a flow of air. The compressed air is then used within the combustor to combust a supply of fuel. The resulting flow of hot gases from the combustion process is expanded through the turbine, which causes the rotor blades to rotate the shaft to which they are attached. In this manner, energy contained in the fuel is converted into the mechanical energy of the rotating shaft, which then, for example, may be used to rotate the coils of a generator to generate electricity.
Referring to FIGS. 1 and 2, turbine rotor blades 100 generally include an airfoil portion or airfoil 102 and a root portion or root 104. The airfoil 102 may be described as having a convex suction face 105 and a concave pressure face 106. The airfoil 102 further may be described as having a leading edge 107, which is the forward edge, and a trailing edge 108, which is the aft edge. The root 104 may be described as having structure (which, as shown, typically includes a dovetail 109) for affixing the blade 100 to the rotor shaft, a platform 110 from which the airfoil 102 extends, and a shank 112, which includes the structure between the dovetail 109 and the platform 110.
The platform 110 may comprise a variety of configurations such as planar (as illustrated in FIG. 4), contoured or any other suitable variations. (Note that “planar,” as used herein, means approximately or substantially in the shape of a plane. For example, one of ordinary skill in the art will appreciate that platforms may be configured to have an outboard surface that is slight curved and convex, with the curvature corresponding to the circumference of the turbine at the radial location of the rotor blades. As used herein, this type of platform shape is deemed planar, as the radius of curvature is sufficiently great to give the platform a flat appearance.) For example, with specific reference to FIG. 4, the platform 110 may have a topside 113, which, as shown in FIG. 1, may include an axially and circumferentially extending flat surface. As shown in FIG. 2, the platform 110 may have a underside 114, which may also include an axially and circumferentially extending flat surface. The topside 113 and the bottom side 114 of the platform 110 may be formed such that each is substantially parallel to the other. As depicted, it will be appreciated that the platform 110 typically has a thin radial profile, i.e., there is a relatively short radial distance between the topside 113 and the bottom side 114 of the platform 110.
In general, the platform 110 may be employed on turbine rotor blades 100 to form the inner flow path boundary of the hot gas path section of the gas turbine. The platform 110 further provides structural support for the airfoil 102. In operation, the rotational velocity of the turbine induces mechanical loading that can create relatively stressed regions along the platform 110 that may further experience elevated temperatures.
One method to make the platform region 110 more durable is to cool it with a flow of compressed air or other coolant during operation. However, as one of ordinary skill in the art will appreciate, the platform region 110 presents certain design challenges. In significant part, this is due to the geometry of this region, in that, as described, the platform 110 is a periphery component that resides away from the central core of the rotor blade and typically is designed to have a structurally sound, but thin radial thickness.
To circulate coolant, rotor blades 100 typically include one or more interior cooling passages 116 (see FIGS. 3 and 4) that, at minimum, extend radially through the core of the blade 100, including through the root 104 and the airfoil 102. As described in more detail below, to increase the exchange of heat, such interior cooling passages 116 may be formed having a serpentine path that winds through the central regions of the blade 100, though other configurations are possible. In operation, a coolant 180 may enter the interior cooling passages 116 via one or more inlets 117 formed in the inboard portion of the root 104. The coolant 180 may circulate through the blade 100 and exit through outlets (not shown) formed on the airfoil and/or via one or more outlets (not shown) formed in the root 104. The coolant 180 may be pressurized, and, for example, may include pressurized air, pressurized air mixed with water, steam, and the like. In many cases, the coolant 180 is compressed air that is diverted from the compressor of the engine, though other sources are possible.
In some cases, the interior cooling passages 116 may further comprise a cavity 119 formed between the shanks 112 and platforms 110 of adjacent rotor blades 100. From there, the coolant 180 may be used to cool the platform region 110 of the blade by connecting cooling passages in the platform to one or more parts of interior cooling passages 116 (e.g. the cavity 119) of the turbine rotor blade. For example, coolant may be extracted from one of the interior cooling passages 116 (such as into the cavity 119 formed between the shanks 112/platforms 110) and supplied to cooling channels that extend through the platforms 110. After traversing the platform 110, the cooling air may exit the cavity through film cooling holes formed in the topside 113 of the platform 110.
It will be appreciated, however, that this type the cooling circuit is not self-contained in one part, as the cooling circuit depending on the cavity 119 is only formed after two neighboring rotor blades 100 are assembled. This may produce greater complexity to installation and testing. An alternative design for platform cooling can comprise a cooling circuit contained within the rotor blade 100 and does not involve the shank cavity 119, as depicted in FIG. 3. Cooling air is extracted from one of the interior cooling passages 116 that extend through the core of the blade 110 and directed aft through cooling channels formed within the platform 110. The coolant 180 flows through the platform cooling channels to help cool the component. However, placing cooling channels in the platform 110 is often achieved in the relatively hottest areas such as on the concave pressure face 106 side of the airfoil 102. Moreover, cooling channel arrangements may be complex (e.g., curved or serpentine) for distribution of coolant 180 thereby requiring casting or other initial manufacturing considerations that prevent or limit making such modifications in pretexting parts.
Accordingly, alternative platform cooling arrangements would be welcome in the art.